1. Technical Field of the Invention
The present invention relates to a cooling structure of a turbine airfoil in a gas turbine for aviation or industry.
2. Description of the Prior Art
In the turbine airfoil of a gas turbine for aviation or industry, since the external surface is exposed to hot gas (e.g., 1000° C. or more) during operation, the turbine airfoil is generally cooled from the inside thereof by flowing cooling gas (e.g., cooling air) into the inside so as to prevent the turbine airfoil from overheating.
In order to improve the cooling performance of the turbine airfoil, several proposals have been suggested (e.g., Patent Documents 1 to 3).
In the gas turbine airfoil disclosed in Patent Document 1, the cooling air is fed from a tube 56 inside an airfoil 50, as shown in FIGS. 1A, 1B and 1C. The cooling air 69 flows toward the internal surface 54 of the airfoil through flow openings 68 of the tube 56. Small, elongated protrusions 61 are installed on at least the same positions as the flow openings 68 of the airfoil internal surface 54. The passage area of a flow passage 58 between the tube 56 and the airfoil internal surface 54 is increased toward an outlet 60 side.
The gas turbine airfoil disclosed in Patent Document 2 includes a first sidewall 70 and a second sidewall 72 which are connected to each other by a leading edge 74 and a trailing edge 76, and a first cavity 77 and a second cavity 78 which are spaced to be separated by a partition wall positioned between the first side wall 70 and the second side wall 72, as shown in FIGS. 2A and 2B. A rearward bridge 80 extends along the first cavity 77, and has a row of outlet holes 84 therein. The partition wall 88 has a row of inlet holes 82. A row of turbulators 86 are arranged on the inside of the first cavity 77, and extend from the first sidewall to the second sidewall. The turbulators 86 are inclined with respect to the inlet holes 82 to perform multiple impingement cooling.
The gas turbine airfoil disclosed in Patent Document 3 includes an external surface 91 facing combustion gas 90 and an internal surface 92 against which cooling gas impinges, as shown in FIG. 3. The internal surface 92 is provided with a plurality of ridges 94 and a plurality of grooves 96 so as to improve heat transfer due to impingement cooling.
Patent Document 1: U.S. Pat. No. 5,352,091 entitled “GAS TURBINE AIRFOIL”
Patent Document 2: U.S. Pat. No. 6,174,134 entitled “MULTIPLE IMPINGEMENT AIRFOIL COOLING”
Patent Document 3: U.S. Pat. No. 6,142,734 entitled “INTERNALLY GROOVED TURBINE WALL”
In general, since the airfoil leading edge of the gas turbine has a large curvature, the cooling side area which comes into contact with the cooling gas is small as compared with the hot side area which is exposed to the high-temperature gas. For this reason, there are many cases where the airfoil leading edge does not obtain the necessary cooling effectiveness only by convection cooling at the cooling sidewall. The turbine airfoil has generally a plurality of film cooling holes through which the cooling air is blown out from the surface of the turbine airfoil, thereby cooling the turbine airfoil by heat absorption at the holes.
Significant quantities of holes are required to cool the turbine airfoil with heat absorption, but if the opening area of the holes is increased, the cooling air is likely to flow backwards at the holes. Therefore, conventionally, the opening area of the impingement holes is increased, and an appropriate pressure difference for the back flow is given. In this instance, however, there is a problem in that the flow rate of the cooling air is increased, so that engine performance deteriorates.